High-Lift Airfoil in Langley Low-Turbulence Pressure Tunnel

158 Pages·2015·16.72 MB·English

https://ntrs.nasa.gov/search.jsp?R=20030013625 2019-04-12T13:18:29+00:00Z
NASA/TM-2002-211780
Experimental Test Results of Energy Efficient
Transport (EET) High-Lift Airfoil in Langley
Low-Turbulence Pressure Tunnel
Harry L. Morgan, Jr.
Langley Research Center, Hampton, Virginia
December 2002
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NASA/TM-2002-211780
Experimental Test Results of Energy Efficient
Transport (EET) High-Lift Airfoil in Langley
Low-Turbulence Pressure Tunnel
Harry L. Morgan, Jr.
Langley Research Center, Hampton, Virginia
National Aeronautics and
Space Administration
Langley Research Center
Hampton, Virginia 23681-2199
December 2002
Available from:
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Summary
An experimental study has been conducted in the Langley Low-Turbulence Pressure Tunnel to
determine the effects of Reynolds number and Mach number on the two-dimensional aerodynamic
performance of the Langley Energy Efficient Transport (EET) High-Lift Airfoil. This high-lift airfoil is a
supercritical-type airfoil with a thickness-to-chord ratio of 0.12 and is equipped with a leading-edge slat
and a double-slotted trailing-edge flap. The two-element trailing-edge flap consisted of a large-chord
vane and small-chord aft flap. All the elements were supported by a set of brackets that held each
element at fixed deflection, gap, and overlap. The leading-edge slat brackets consisted of a set of four
brackets with deflections of-30 °,-40 °,-50 °,and -60 °. The trailing-edge flap brackets were designed for
equal deflections between the main and vane elements and between the vane and aft-flap elements and
consisted of a set of four brackets with deflections of 7.5°, 15°, 22.5°, and 30°. These sets of slat and flap
brackets resulted in 16different configurations each with accurately defined and highly repeatable lofted
geometries. The model was equipped with a densely defined row of chordwise surface pressure taps
along the model midspan and two coarsely defined chordwise rows 2.5 in. from each sidewall. The
aerodynamic forces and moment were measured by a yoke-type three-component, strain-gauge balance
and model support system that had an angle-of-attack range of-8 °to 26°. All 16 configurations were
tested at a free-stream Mach number of 0.20 and, for a few selected configurations, through a Mach
number range of 0.10 to 0.35. In addition, all of the configurations were tested through a Reynolds
number range of 2.5 x 106to 18 x 106. For a few selected configurations, the drag was measured with a
downstream mounted wake traversing system that held a rake consisting of three equally spaced, five-
hole pressure probes. During the testing, the spanwise two-dimensionality of the flow over the model
was controlled by energizing the tunnel sidewall boundary layer to delay or prevent separation with a set
of four tangential blowing slots located at specific locations on each model endplate.
The test results demonstrate the tremendous effect of Reynolds number and Mach number on the
aerodynamic performance of this supercritical-type high-lift airfoil. Analysis of the test data revealed
several inconsistencies inthe trends observed showing the effects of an increase in Reynolds number and
Mach number on the maximum lift performance of the high-lift airfoil. The endplate blowing system
developed was able to adequately control the separation of the sidewall boundary layer; thereby, spanwise
uniformity of the flow around the model during the test was maintained. The model geometry, surface
pressures, balance-measured forces and moment, and wake data obtained are very well defined for all 16
configurations tested; therefore, these data are well suited for the validation and calibration of computer
codes that predict high-lift system performance and flow field characteristics.
Introduction
During the early 1970s through the late 1980s the National Aeronautics and Space Administration
was actively involved in an aeronautical research effort to improve the energy efficiency of modern wide-
body jet transport aircraft. The Aircraft Energy Efficiency (ACEE) proj ect was formulated to encourage
industry participation and to coordinate the industry and NASA research efforts. One element of the
ACEE project was the Energy Efficient Transport (EET) program, which was concerned primarily with
the development of advanced aerodynamics and active-controls technology for application to derivative
or next-generation transport aircraft. A part of the EET program was the development, by NASA Langley
Research Center personnel, of advanced supercritical wings with greater section thickness-to-chord ratios,
higher aspect ratios, higher cruise lift coefficients, and lower sweepback than the conventional wings of
current transports. These supercritical wings were tested extensively inthe Langley wind tunnels to
determine their high-speed cruise performance characteristics. Because of their high cruise lift
coefficients and high aspect ratios, these wings could be smaller and more fuel efficient than wings used
currently provided high-lift flaps systems could be designed to ensure that takeoff and landing
requirements could be met.
As part of the EET Program, a high-lift flap system was designed for a representative supercritical
wing and tested on both a two-dimensional airfoil model and on two different scaled three-dimensional
wing models. One high-lift wing model with a span of 7.5 fl was tested at high Reynolds number, high-
pressure conditions in the Ames 12-Foot Pressure Tunnel. The other model with a span of 12 fl was
tested at low Reynolds number, atmospheric conditions in the Langley 14- by 22-Foot Subsonic Tunnel.
The 7.5-fl span model was also tested in the Langley 14- by 22-Foot Subsonic Tunnel to obtain support
system interference and wall corrections for the Ames tests. Both models had an aspect ratio of 12, a
quarter-chord sweep of 27°, and the wing and body shape of the NASA supercritical SCW-2a high-speed
transonic model tested in the Langley 8-Foot Transonic Pressure Tunnel and reported in references 1and
2. Both high-lift models were tested extensively from the late 1970s to the mid 1980s and the data are
reported in references 3 through 9. A photograph of the 12-fl span model mounted in the Langley 14- by
22-Foot Subsonic Tunnel is shown in figure 1.
The high-lift flap for these models consisted of a part-span double-slotted trailing-edge flap and a
full-span leading-edge slat. The trailing-edge flap consisted of a large-chord vane and small-chord aft flap
combination, as opposed to the more conventionally used small-chord vane and large-chord aft flap
combinations. Vane-flap combinations similar to the combination used on these models had also been
under development by several aircraft manufacturers and had achieved maximum two-dimensional lift
coefficients approaching those of more complex triple-slotted flap combinations. Each model was also
equipped with inboard high-speed ailerons, outboard low-speed ailerons, two wing-mounted flow-through
nacelles, landing gear, movable horizontal tails, and interchangeable wingtips that provided for aspect
ratios of both 10 and 12. Each model was instrumented with a six-component strain-gauge balance to
measure aerodynamic forces and moments and with chordwise pressure taps at three spanwise stations to
determine representative wing and flap loads.
The cruise wing for these three-dimensional high-lift models had a break station at the 38.3-percent
semispan location as shown in figure 2. The airfoil t/c at this location is 0.12 and was close to the average
t/c of the wing, which has a root t/c of 0.144 and a tip t/c of 0.10. The high-lift flap system for the wing
was designed first by defining the element shapes at the break station and then extending those shapes to
the inboard and outboard wing location through linear extrapolation. A constant-chord model of the high-
lift airfoil at the wing break station was built and tested in the Langley Low-Turbulence Pressure Tunnel
(LTPT). The results from the test of that model are presented in this report. These data cover a range of
Reynolds numbers from 2.5 x 106to 18 x 106and Mach numbers from 0.10 to 0.35. The data consist of
chordwise surface static pressures on each element and tunnel centefline floor and ceiling pressures from
the tunnel pressure scanning system, section lift and pitching-moment data from the tunnel balance
system, and selective drag data from the tunnel wake rake survey system.
Symbols
Af balance measured axial force, lb
AR aspect ratio of EET High-Lift Wing Model
b model span, 36.0 in.
Cp local surface static pressure coefficient
airfoil reference chord, 21.654 in.
Cd section drag coefficient
t wake point drag coefficient
Cd
section lift coefficient
C1
section pitching-moment coefficient
Gin
d,,_ distance from model weight center to endplate center of rotation, in.
H_r sidewall blowing-box thrust distance from center of turntable (positive up), in.
h tunnel height, 90.0 in.
tit wake probe height, in.
M free-stream Mach number
m blowing-box mass flow, slugs/min
Nf balance measured normal force, lb
balance measured pitching moment, in-lb
Pin
q_ free-stream dynamic pressure, lb/in2
Rn Reynolds number based on reference chord
Tbx sidewall blowing-box thrust, lb
t/c airfoil thickness-to-chord ratio, 0.12
W_ model weight, lb
x distance along chord of model, in.
Y distance perpendicular to chord ofmodel, in.
z distance along span of model, in.
angle of attack (positive nose up), deg
A deflection angle between longest chords of adjacent elements, deg
8 slat, vane, or flap deflection (positive for trailing edge down), deg
8s solid blockage correction factor
8w wake blockage correction factor
n nondimensional spanwise position, z/b
body shape correction factor
wall correction factor
13
qb sidewall blowing-box thrust angle, deg
deflection of element longest chord, deg
Subscripts:
bx blowing box
corrected
C
f flap
le (L.E.) leading edge
lg longest chord
max maximum
ps wake static pressure
pt wake total pressure
S slat
te (T.E.) trailing edge
uncorrected
u
v vane
Wind Tunnel and Test Apparatus
Wind Tunnel
The EET High-Lift Airfoil test was conducted inthe Langley Low-Turbulence Pressure Tunnel
(LTPT). The LTPT is a single-return, closed-throat wind tunnel that can be operated at tunnel total
pressures from near vacuum to 10atmospheres (ref 10). A sketch of the tunnel circuit arrangement is
shown in figure 3. The tunnel test section is 3 ft wide, 7.5 ft high, and 7.5 ft long, which when combined
with a 17.6-to-1 contraction ratio makes the LTPT ideally suited for low-turbulence, two-dimensional
airfoil testing. The Reynolds number capability of the tunnel for a typical high-lift airfoil test is shown in
figure 4. The tunnel can achieve amaximum Reynolds number of 15 x 106per foot ataMach number of
0.24. The maximum empty-tunnel speed at a total pressure of 1atmosphere is a Mach number of 0.47
with a corresponding Reynolds number of 3 x 106per foot. The tunnel total temperature iscontrolled
through a set of internal heat exchange coils located upstream of the screens inthe contraction section of
the tunnel. During the warmer months of operation, cooling water is pumped through the heat exchanger
and circulated through the cooling tower located in the inner courtyard. During the colder months of
operation, the circulation water is heated by a stream injection system.
Model-Support and Force-Balance System
During the early 1970s anew model-support and force-balance system capable of handling both
single-element and multielement airfoils was installed inthe LTPT to provide the capability for two-
dimensional high Reynolds number testing. A sketch of this model-support and force-balance system is
shown in figure 5. An airfoil model is mounted between two endplates that are connected to the inner
drums. These inner drums are held in place by an outer drum and yoke arm support system. The yoke
arm support system is mounted to the force balance, which is connected to the tunnel through a balance
platform. The attitude of the model is controlled by a motor-driven, externally mounted pitch mechanism
that rotates the bearing-mounted inner drums. A multipath labyrinth seal is used to minimize air leakage
from the test section into the outer tunnel plenum.
The force balance isathree-component strain-gauge balance of the external virtual-image type.
The maximum balance loads are 18000 lb in lift, 550 lbin drag, and 12000 ft-lb inpitching moment. The
balance istemperature compensated and calibrated to account for first- and second-order interactions, and
ithas a general accuracy of+0.5 percent of design loads.
Sidewall Boundary-Layer Control System
To ensure spanwise uniformity of the flow field when testing high-lift airfoils near the maximum
lift condition, some form of tunnel sidewall boundary-layer control (BLC) was needed. The large adverse
pressure gradients induced on the tunnel sidewalls by ahigh-lift airfoil near maximum lift can cause the
sidewall boundary layer to separate with a corresponding loss of spanwise uniformity of the flow on the
airfoil surface and aresulting premature loss of lift. Because asource of high-pressure air was available
for the LTPT, tangential blowing was selected as the means of providing sidewall BLC during the tests of
this high-lift airfoil. Four blowing boxes with tangential blowing slots were mounted on the model
endplates on both sides of the tunnel and were positioned around the airfoil within the confines of the
endplates. High-pressure air was supplied to each box through a flexible hose connected to the blowing-
box control cart with remote-controlled valves for each box. A cross-sectional sketch of a typical blowing
box ispresented infigure 6. The blowing boxes were designed to provide uniform tangential flow at the
slot exit. High-pressure air flows into an inner manifold distribution chamber and is then distributed
through slots to an outer manifold chamber. An adjustable slot lip and the box itself form the exit slot.
For this test, the width of the slot exit for all the boxes was set at 0.060 in. and the box supply air
pressures were adjusted to achieve the maximum mass-flow rate through the boxes. The chordwise
locationandslotlengthsforeachofthefourboxesarepresenteindfigure7.
Thetangentiafllowofairfromtheblowingboxesontheendplatepsroduceadthrustingforceand
skin-frictionforceintheupstreamdirectionthatwasconsidereadtareloadontheforce-balanscyestem.
Duringtheinitialpartofthetest,wind-offtarerunswereperformeadtdifferenttunnetlotalpressures
withtheboxmassflowssetatmaximumT.hesedatawerecurvefit andasetoftarevaluesderivedthat
weresubtractefrdomthemeasurewdind-ondata.
Remote-Controlled Wake Survey Apparatus
A limited amount of airfoil drag data was computed during this investigation with the momentum
method applied to the measured downstream wake properties. The momentum deficient in the wake was
measured with a pressure probe that was traversed through the wake by a remotely controlled traverse
system. Detailed descriptions of the mechanism, the calibration of the probe head, and the drag equations
used are given inreference 10. A sketch of the wake traverse apparatus ispresented in figure 8. The
vertical support strut attaches the wake rake assembly tothe tunnel sting-support arc sector and houses the
traverse system. The wake traverse system provides vertical motion of the pressure rake within a total
range of 47 in. The vertical drive mechanism consists of a vertically mounted direct-current stepper
motor that drives a ball screw, which, in turn, drives the exterior traverse arm. An optical shaft encoder
tracks the vertical position with a position accuracy of 0.0005 in. The probe head is attached to the pitch
arm, which is supported by the exterior traverse. Extension arms can be placed between the exterior
traverse and the pitch arm to provide the capability to position the probes at streamwise locations of 22
in., 33 in., and 44 in. downstream from the turntable center of rotation.
During this investigation, the probes were positioned at the 44-in. location. A sketch of the pitch
arm, probe head, and pressure probes isshown in figure 9. The probe head can be pitched about its pitch
arm attachment point within a + 45°range. This motion is driven by a pitch link mechanism that is
controlled by a globe gear motor. The probe head also has a variable roll orientation capability. The
probe head tip rotates relative to a fixed inner cylinder that can be locked into several roll angle positions.
These fixed roll positions are 0°, 7.6 o,30o,48.6 o,and 90oas indicated on the cross-sectional drawings of
the probe head in figure 9. This particular set of roll angles provides the capability to take spanwise
measurements at 0, 0.5, 1, 2, 3, 4, 6, and 8in. from the centerline. The roll axis of the probe head is
located at the midspan location (18 in.) of the tunnel test section. During this investigation, the roll
orientation was set at 90o,which placed probe 1at the model midspan, probe 2 at 4 in. off the midspan,
and probe 3 at 8 in. off the midspan. A photograph of the probe head rotated to the 0oposition is shown
in figure 10.
High-Lift Airfoil Model
The high-lift airfoil tested during this investigation has been designated as the Langley EET High-
Lift Airfoil. The cruise airfoil with all elements nested has the same coordinates as those of the wing
section at the break station of the NASA supercritical SCW-2a wing described in references 1and 2. This
high-lift airfoil has a thickness-to-chord ratio of 0.12 and a chord of 21.654 in. Normally airfoils built for
testing inthe LTPT have a chord of 24 in.; however, after the high-lift flap system was designed and
deflections of the elements set, itwas found that a slightly reduced chord would ensure that all the
deflected elements would fit within the contours of the endplates. This was important because the tunnel
walls start to diverge just downstream of the aft edge of the wall endplates; therefore, any aft element
surfaces that extended beyond the endplate would produce a gap between the element edge and the wall
that would require the addition of filler material. The EET High-Lift Airfoil had a span of 36 in. It was
designed to operate at the maximum tunnel operating conditions of 10 atmospheres at a Mach number of
0.2. The airfoil had an area of 5.414 ft2,and at the maximum tunnel dynamic pressure of 576 lb/ft2and
with an anticipated maximum lift coefficient of 4.5, the resultant lift force would be approximately 14000